The present invention relates generally to gas turbine engines, and, more specifically, to turbine blade cooling therein.
In a gas turbine engine air is pressurized in a compressor and mixed with fuel in a combustor for generating hot combustion gases. Turbines are used to extract energy from the gases and power the compressor while producing useful output power such as driving an upstream fan in an aircraft turbofan gas turbine engine application.
Engine efficiency may be maximized by maximizing the temperature of the combustion gases, but the high combustion gas temperature will limit the useful life of the various turbine components exposed to the combustion gases during operation.
The first stage turbine rotor blade receives the hottest combustion gases from the upstream turbine nozzle in the high pressure turbine (HPT). These blades have dovetails mounted in corresponding dovetail slots in the perimeter of a supporting rotor disk, and airfoils extend outwardly from a flow boundary platform mounted on the dovetails.
The turbine airfoils are hollow and include various internal cooling circuits therein having respective inlets extending through the platform and dovetail for receiving cooling air from the base of the dovetail mounted in the dovetail slots. The cooling air is typically compressor discharge air having maximum pressure, along with maximum temperature due to the compression process.
The typical operating cycle for an aircraft turbofan engine includes takeoff, climb, cruise, descent, and landing during which thrust reverse operation is temporarily effected. Maximum power operation of the engine is typically effected during takeoff during which the turbine rotor inlet temperature may reach a corresponding maximum value, along with a corresponding maximum temperature for the compressor discharge air.
The cooling circuits for the first stage turbine blades may therefore be designed for this maximum temperature condition during takeoff, which condition is transient and of relatively short duration.
Accordingly, state-of-the-art superalloy materials, typically nickel or cobalt based, are used in the casting of the first stage turbine rotor blades for maximizing their strength at elevated temperature and ensuring their durability and long useful life. Correspondingly, the airfoil cooling circuits may be configured in a myriad of permutations for maximizing the cooling effectiveness of the hot compressor discharge air in the different regions of the airfoil subject to different heating loads from the combustion gases which flow with different pressure and temperature distributions around the opposite pressure and suction. sides of the airfoil.
The compressor discharge air typically used for cooling the airfoil is initially channeled inside the hollow airfoil and is then discharged through various rows of aperture outlets in the pressure and suction sides thereof. The compressor discharge air has maximum pressure and is used to ensure a suitable backflow margin at the various outlets in the turbine airfoils. The combustion gases decrease in pressure as they flow downstream over the leading and trailing edges of the airfoils, and sufficient backflow margin must be provided along the airfoil leading edge wherein the local pressure of the combustion gases is relatively high.
A typical backflow margin requires that the pressure of the cooling air in the airfoil exceed the local pressure of the combustion gases outside thereof by about five to fifty percent. In this way, the combustion gases are not back-ingested into the airfoil through the outlets for maintaining proper cooling effectiveness of the internal circuits.
As the combustion gases decrease in pressure to the trailing edge of the airfoil, the local backflow margin correspondingly increases due to the relatively high pressure of the compressor discharge air channeled into the airfoils. Excess backflow margin is not desirable since it leads to blow-off or lift-off of the spent cooling air as it is discharged from the outlet holes in typical film cooling configurations.
The airfoil internal cooling circuits are therefore typically tailored for the different operating conditions between the leading and trailing edges of the airfoil. The leading edge cooling circuit typically provides internal impingement cooling of the back side of the leading edge followed by discharge of the spent impingement air through various rows of film cooling holes around the airfoil leading edge.
The trailing edge cooling circuit typically includes either centerline or pressure-side outlet holes along the trailing edge fed from an internal radial channel. The middle or mid-chord region of the airfoil typically includes a multi-pass serpentine circuit having radial legs through which the cooling air is channeled and absorbs heat prior to discharge through various outlet apertures.
The various internal cooling circuits typically include elongate turbulators or ribs extending along the pressure and suction sidewalls of the airfoil for increasing the heat transfer capability of the cooling air. The turbulators and specific configurations of the cooling circuits introduce pressure losses or pressure drops in the cooling air prior to discharge from the various outlets.
In an advanced turbofan gas turbine engine being developed for small commercial business jets or military applications, the core engine is being designed to operate substantially continuously at very high compressor discharge temperature and correspondingly high turbine rotor inlet temperature for extended periods of time. In contrast with conventional turbofan engines having turbine blades designed for transient takeoff temperature conditions, the advanced turbofan engine requires turbine cooling configurations designed for long duration high temperature conditions.
Accordingly, the turbine blades require a substantially lower bulk temperature during normal operation than required for typical turbofan engines. The requirement for lower bulk temperature of the turbine airfoils therefore requires improved cooling circuits which better maximize the cooling effectiveness of the correspondingly high temperature compressor discharge air.
It is therefore desired to provide a turbine blade having an improved cooling configuration therein for effecting a lower bulk temperature during operation.